This invention relates to the field of gas turbine engines and, in particular, to gas turbine engines having a can annular combustor.
Gas turbine engines are known to include a compressor for compressing air; a combustor for producing a hot gas by burning fuel in the presence of the compressed air produced by the compressor, and a turbine for expanding the hot gas to extract shaft power. The combustion process in many older gas turbine engines is dominated by diffusion flames burning at or near stoichiometric conditions with flame temperatures exceeding 3,000xc2x0 F. Such combustion will produce a high level of oxides of nitrogen (NOx). Current emissions regulations have greatly reduced the allowable levels of NOx emissions. Lean premixed combustion has been developed to reduce the peak flame temperatures and to correspondingly reduce the production of NOx in gas turbine engines. In a premixed combustion process, fuel and air are premixed in a premixing section of the combustor. The fuel-air mixture is then introduced into a combustion chamber where it is burned. U.S. Pat. No. 6,082,111 describes a gas turbine engine utilizing a can annular premix combustor design. Multiple premixers are positioned in a ring to provide a premixed fuel/air mixture to a combustion chamber. A pilot fuel nozzle is located at the center of the ring to provide a flow of pilot fuel to the combustion chamber.
The design of a gas turbine combustor is complicated by the necessity for the gas turbine engine to operate reliably with a low level of emissions at a variety of power levels. High power operation at high firing temperatures tends to increase the generation of oxides of nitrogen. Low power operation at lower combustion temperatures tends to increase the generation of carbon monoxide and unburned hydrocarbons due to incomplete combustion of the fuel. Under all operating conditions, it is important to ensure the stability of the flame to avoid unexpected flameout, damaging levels of acoustic vibration, and damaging flashback of the flame from the combustion chamber into the fuel premix section of the combustor. A relatively rich fuel/air mixture will improve the stability of the combustion process but will have an adverse affect on the level of emissions. A careful balance must be achieved among these various constraints in order to provide a reliable machine capable of satisfying very strict modern emissions regulations.
Dynamics concerns vary among the different types of combustor designs. Gas turbines having an annular combustion chamber include a plurality of burners disposed in one or more concentric rings for providing fuel into a single toroidal annulus. U.S. Pat. No. 5,400,587 describes one such annular combustion chamber design. Annular combustion chamber dynamics are generally dominated by circumferential pressure pulsation modes between the plurality of burners. In contrast, gas turbines having can annular combustion chambers include a plurality of individual can combustors wherein the combustion process in each can is relatively isolated from interaction with the combustion process of adjacent cans. Can annular combustion chamber dynamics are generally dominated by axial pressure pulsation modes within the individual cans.
Staging is the delivery of fuel to the combustion chamber through at least two separately controllable fuel supply systems or stages including separate fuel nozzles or sets of fuel nozzles. As the power level of the machine is increased, the amount of fuel supplied through each stage is increased to achieve a desired power level. A two-stage can annular combustor is described in U.S. Pat. No. 4,265,085. The combustor of the ""085 patent includes a primary stage delivering fuel to a central region of the combustion chamber and a secondary stage delivering fuel to an annular region of the combustion chamber surrounding the central region. The primary stage is a fuel-rich core wherein stoichiometry can be optimized. U.S. Pat. No. 5,974,781 describes an axially staged hybrid can-annular combustor wherein the premixers for two stages are positioned at different axial locations along the axial flow path of the combustion air. U.S. Pat. No. 5,307,621 describes a method of controlling combustion using an asymmetric whirl combustion pattern.
With the continuing demand for gas turbine engines having lower levels of emissions and increased operational flexibility, further improvements in gas turbine combustor design and operation are needed. Accordingly, a can combustor for a gas turbine engine is described herein as including: a first stage comprising a first plurality of burners arranged symmetrically around a longitudinal centerline of a combustion chamber at a first radial distance from the centerline; and a second stage comprising a second plurality of burners arranged symmetrically around the centerline of the combustion chamber at a second radial distance different than the first radial distance. The burners of the second stage may be angularly positioned midway between respective neighboring burners of the first stage or at respective angular locations other than midway between respective neighboring burners of the first stage.
A can combustor for a gas turbine engine is further describe as including: a first stage comprising a first plurality of burners arranged symmetrically around a longitudinal centerline of a combustion chamber and angularly separated from each other by an angle of 360/N degrees; a second stage comprising a second plurality of burners arranged symmetrically around the longitudinal centerline of the combustion chamber and angularly separated from each other by an angle of 360/N degrees; wherein the burners of the second stage are positioned at respective angular locations other than midway between respective neighboring burners of the first stage. The first plurality of burners may be spaced from the longitudinal centerline at a first radial distance; and the second plurality of burners may be spaced from the longitudinal centerline at a second radial distance different than the first radial distance.
A gas turbine engine is described as including: a compressor for supplying compressed air; a can annular combustor for burning fuel in the compressed air to produce a hot gas; and a turbine for expanding the hot gas; wherein the can annular combustor further comprises a plurality of can combustors each comprising: an annular member defining a combustion chamber having a longitudinal centerline; a first plurality of burners disposed in a symmetrical ring around the centerline at a first radial distance; and a second plurality of burners disposed in a symmetrical ring around the centerline at a second radial distance greater than the first radial distance. The angular position of the second plurality of burners may be selected so that the burners of the second plurality of burners are angularly centered between respective neighboring burners of the first plurality of burners or so that the burners of the second plurality of burners are not angularly centered between respective neighboring burners of the first plurality of burners.
A gas turbine engine is describe herein as including: a compressor for supplying compressed air; a can annular combustor for burning fuel in the compressed air to produce a hot gas; and a turbine for expanding the hot gas; wherein the can annular combustor further comprises a plurality of can combustors each comprising: a first stage of burners disposed in a symmetrical circular pattern about a centerline, N being the number of burners in the first stage of burners and 360/Nxc2x0 being an angle of separation between burners of the first stage of burners; a second stage of burners disposed in a symmetrical circular pattern about the centerline, the burners of the second stage of burners being singularly disposed between respective neighboring burners of the first stage of burners, N being the number of burners in the second stage of burners and 360/Nxc2x0 being an angle of separation between burners of the second stage of burners; and an angular separation between burners of the first stage of burners and neighboring burners of the second stage of burners being an angle not equal to 360/2Nxc2x0. The first stage of burners may be disposed in a circular pattern having a first radius about the centerline; and the second stage of burners may be disposed in a circular pattern having a second radius about the centerline not equal to the first radius.